1. Field of the Invention
The present invention relates to an attitude determination system for an artificial satellite for determining attitude of a spacecraft or an artificial satellite on the basis of identification as to whether the stars on an all-sky star catalog are caught by a star sensor. Incidentally, the term "artificial satellite" is used herein as being equivalent to the term "spacecraft".
2. Description of Related Art
For having better understanding of the concept underlying the present invention, description will first be made in some detail of a conventional artificial satellite attitude determination system. FIG. 13 is a schematic diagram for illustrating a star identifying operation in a conventional artificial satellite attitude determination system which is disclosed, for example, in Japanese Unexamined Patent Application Publication No. 6100/1985 (JP-A-60-6100). Referring to the figure, reference numeral 1 designates an artificial satellite equipped with a star sensor (not shown) and traveling or cruising around the earth 2, numeral 3 designates a plurality of stars on the celestial sphere observed by the star sensor mounted on the artificial satellite 1, numeral 4 designates a star concerned of those observed by the star sensor, numeral 5 designates a circle having a radius corresponding to a maximum value of the angle of field of the star sensor, numeral 6 designates all the stars that are darker than the star 4 concerned within the circle 5, numeral 7 designates a set including collected star patterns in all of the sky, numeral 8 designates a table containing combinations of the star patterns in all of the sky and all characteristic quantities, respectively, numeral 9 denotes a table prepared by rearranging the table 8 in the descending order of the characteristic quantities, and numeral 10 denotes a table generated by extracting only the pattern identifying numbers from the table 9. A star catalog can be created on the basis of the contents of the tables mentioned above.
Next, description will turn to the operation of the conventional artificial satellite attitude determination system. The artificial satellite 1 carrying the star sensor travels around the earth 2 while observing a plurality of stars 3 on the celestial sphere by means of the star sensor. The characteristic quantities required for the star identification are determined in such a manner as mentioned below. For all the stars on the celestial sphere which can be caught by the star sensor, the circle 5 is depicted around the concerned star 4 assumed as being positioned at the center with a radius corresponding to the maximum value of the angle of field of the star sensor, whereon the stars 6 falling within the circle and darker than the central star 4 are sorted out. By drawing lines radially from the concerned star 4 located at the center to the individual darker stars 6, a radial line pattern is generated. The star identifying number of the concerned star 4 is affixed to the generated radial line pattern as the pattern identifying number.
On the basis of the set 7 of the patterns for the stars in all the sky, the table 8 is generated by combining the characteristic quantities and the pattern identifying numbers of all the patterns. In the table 8, the characteristic quantities and the pattern identifying numbers are rearranged in the descending order of the magnitudes of the characteristic quantities, whereby the table 9 is generated in which the characteristic quantities and the pattern identifying numbers are listed in the descending order of the characteristic quantities. On the basis of the table addresses of the table 9, approximating functions 11 for representing the characteristic quantities are generated. Additionally, the pattern identifying number table 10 containing only the pattern identifying numbers extracted from the table 9 is created. For the star identification, the pattern generated through the similar procedure for the star which is brightest in the image data obtained through observation by the star sensor and the functions 11 are matched or collated with each other, to thereby extract a common or shared portion from a corresponding sub-table of the pattern identifying number table 10. When only one shared portion is extracted, this means that the star identification has been carried out successfully.
FIG. 14 is a block diagram showing a typical configuration of a conventional artificial satellite attitude determination system equipped with a star sensor. This artificial satellite attitude determination system is comprised of an attitude propagation unit 12, an attitude updating unit 13 and a time-independent drift estimation module 13a. Heretofore, in the system for determining the attitude of the artificial satellite with the aid of combination of a star sensor 16 and a gyro 14, it has been a general practice to estimate the drift of the gyro 14 with the precise attitude angle determined on the basis of the output of the star sensor 16. More specifically, the output or observation values of the star sensor 16 are arithmetically processed by a star sensor processing unit 15 to determine an attitude angle qm of the artificial satellite, whereon a difference qe between the attitude angle qm and an estimated attitude angle qh obtained as the output of the attitude propagation unit 12 is inputted to the attitude updating unit 13 as an estimation error.
In the attitude updating unit 13, the drift .omega.d of the gyro 14 is estimated through the medium of a first-order filter KPO+KIO/s on the presumption that the drift is time-invariable, whereon the estimated value .omega.d is subtracted from the gyro output .omega.m to obtain an estimated value .omega.h of the attitude angular velocity (also known as the attitude rate) of the satellite. The attitude propagation unit 12 is so designed as to add the estimated value of the attitude angle of the artificial satellite one sampling cycle before by integrating the estimated value .omega.h of the attitude angular velocity as a function of time and output additionally an estimated value qh of the attitude angle at the current time point.
FIG. 15 is a view for graphically illustrating steady characteristics of attitude angle estimation errors of a navigation filter employed in the conventional attitude determination system. As can be seen in the figure, in the case where the nominal angular velocity (or nominal attitude rate) of the artificial satellite assumes a constant value, the error involved in estimating the attitude angle of the satellite is substantially equal to zero, indicating the desirable characteristics. However, when the nominal angular velocity changes periodically as in the case of the artificial satellite in which steering operation is performed about a yaw axis, error in the attitude angle estimation makes appearance significantly in dependence on the periodical change of the nominal angular velocity.
In the conventional star identifying method known heretofore, identification of the star is performed by collating the characteristic quantities arithmetically determined on the basis of the star images falling within the visual field of the star sensor at a given time point with the stars on the star catalog. Consequently, the conventional star identifying method is imposed with requirement that the visual field of the star sensor has to be set widely so that a sufficient number of stars required for the identification make appearance within the visual field without fail or the identification has to be given up at a time point when a sufficient number of stars can not be seen within the visual field.
As the characteristic quantities mentioned above, there are selected such quantities which remain invariable independent of rotation about the optical axis of camera such as the elongation, i.e., an angle formed by the direction vectors of two stars, an area of a triangle constituted by three stars on an imaging screen of the camera. Thus, the identification which makes much of a relative relation between the predicted star position and the observed star position has not always been realized when the positions of the stars on the imaging screen of the camera can roughly be predicted from the star catalog as in the case where the attitude value of the artificial satellite can be estimated.
Additionally, when the identification is to be performed by a computer system mounted on the artificial satellite, it is necessarily required that the direction of the sight of the star sensor is known roughly in advance and that the range on the star catalog to be searched is sufficiently limited. Unless such computational capability can be ensured by the satellite-onboard computer, it is necessary to send the observation data to a ground station where the arithmetic operations for the identification are carried out, the result of which is then sent back to the artificial satellite.
On the other hand, when the attitude determination of the artificial satellite is performed on the basis of the result of the identification, it is required to determine parameters representing the coordinate transformation or rotational motion between the star catalog data concerning the direction vectors of the stars in the inertial space and the direction vectors of the stars in the coordinate system fixed to the body of the artificial satellite (hereinafter referred to as the spacecraft body coordinate system) which is calculated on the basis of the positions of the star images within the visual field of the star sensor.
Heretofore, it has been a conventional practice to determine at first roughly the attitude of the artificial satellite by using the attitude sensor such as an earth sensor or sun sensor and thereafter determine the attitude of the artificial satellite with high accuracy by resorting to the use of the star sensor, which makes it possible to utilize the rough values with enhanced accuracy for determining the parameters expressing the coordinate transformation or rotational motion mentioned previously. However, in the case of the artificial satellite which is equipped with neither the earth sensor nor the sun sensor, or when these sensors mounted on the artificial satellite suffer fault and can not be used, the rough value mentioned above can not be made use of in capturing initially the attitude of the artificial satellite. To cope with such situation, some measures have to be provided for enabling such rough value to be utilized.
On the other hand, when the rough value such as mentioned above is available, it is conventional practice to reduce overhead involved in the calculations required for estimation of the attitude angle by adopting linearization based on the minute-angle approximation concept. However, when the accuracy of the rough value is low or poor, error due to the approximation can no more be neglected, incurring degradation in convergence in the transient response, giving rise to a problem.
Furthermore, the artificial satellite destined for astronomical and/or terrestrial observation is required to be equipped with the attitude determination system capable of correcting drifts of a gyro with very high accuracy by utilizing the highly accurate attitude angle information available from the star sensor. However, in the conventional attitude determination system known heretofore, the drift of the gyro is regarded as being a constant or a variable of very large time constant on the presumption that the nominal angular velocity or attitude rate of the artificial satellite is constant. As a consequence, when the nominal angular velocity of the artificial satellite changes as a function of time lapse, then the time-dependent variable component of the drift in the gyro can not be estimated with sufficiently high accuracy. Thus, the attitude determination accuracy as realized is not satisfactory.